Gas turbine engine airfoils, such as high pressure turbine vanes, are typically cooled by compressor bleed air. Conventional turbine vanes, such as the one shown at 9 in FIG. 1, generally have a radially inner band or platform 11 and a plenum 13 defined below the platform 11 for receiving the compressor bleed air. Film cooling holes 15 typically extend from the underside of the platform 11 to the platform radially outer surface 17 (i.e. the platform surface facing the hot gas stream). The air flowing from the holes 15 forms a thin cooling film on the radially outer surface 17 of the platform 11.
One disadvantage of the above vane cooling scheme is that it requires additional cooling air to purge the turbine cavity between the adjacent rows of vanes and turbine blades. Furthermore, the film cooling holes must be sufficiently long to allow the cooling air to flow from the plenum to the gas path side of the platform, which results in greater turbine vane manufacturing costs.